Planetary gearbox

ABSTRACT

The invention concerns a planetary gear box with a flow guidance device for a planet gear lubricant stream in the planetary gear box, which is directed at least partially along a planet gear after an engagement with a sun gear, characterized by a collecting region of the flow guidance device in which the planet gear lubricant stream between the planet gear and the flow guidance device can be at least partially captured, and wherein the collecting region has an outlet region between the flow guidance device and the tooth face of the planet gear for at least a part of the planet gear lubricant stream, and an edge is arranged at the outlet region for defined constriction of the passage for the planet gear lubricant stream.

This application claims priority to German Patent Application102021209552.6 filed Aug. 31, 2021, the entirety of which isincorporated by reference herein.

The present disclosure relates to a planetary gear box having featuresas disclosed herein.

The lubrication and cooling of planetary gear boxes, usually by means ofoil, is of great importance since there are numerous tooth flankcontacts inside the gear box. In particular when used in gas turbineengines of aircraft, lubrication of great importance since theseplanetary gear boxes must remain functioning for a very long timewithout the possibility of maintenance of the planetary gear boxes.

There is therefore a need to provide planetary gear boxes with anefficient lubrication.

According to a first aspect, a planetary gear box has a flow guidancedevice for a planet gear lubricant stream in the planetary gear box. Theplanet gear lubricant stream is a lubricant stream which forms in theplanetary gear box and extends along the tooth surfaces and/or theperiphery of a planet gear.

The planet gear lubricant stream is oriented at least partially along aplanet gear, in particular after an engagement with a sun gear.

The flow guidance device has a collecting region in which the planetgear lubricant stream between the planet gear and the flow guidancedevice can be at least partially captured, and wherein the collectingregion has an outlet region between the flow guidance device and thetooth face of the planet gear for at least a part of the planet gearlubricant stream. This means that a part of the planet gear lubricantstream is conveyed through the outlet region, i.e. oil is present forexample between the tooth flanks and/or in the gap at the planet gear,while another part of the oil may e.g. be discharged.

An edge is arranged at the outlet region for defined constriction of thepassage for the planet gear lubricant stream.

By use of a sharp edge, the proportion of the lubricant carried along bythe rotation of the planet gear may be reduced so that, in particular,crushing losses of oil on the ring gear can be avoided. Thus an ageingof the oil can be avoided, leading to a longer service life.

In one embodiment, an angle, measured in the rotational direction of theplanet gear between the edge and a tangent to the planet gear, isgreater than 90°, in particular less than 180°, quite particularly lessthan 120°. Thus the planet gear lubricant stream flowing onto the edgealways meets a hard edge.

Also, the main flow direction of the planet gear lubricant stream may beoriented from an oil supply to the collecting region of the flowguidance device, the edge or the outlet region of the flow guidancedevice, so that as much lubricant as possible is captured.

Since a geometrically exact alignment of the main flow direction of theplanet stream may not always be possible with accuracy, the main flowdirection may deviate by +/−20°, in particular +/−10° from an exactlinear geometric orientation.

For an efficient lubricant supply, in one embodiment, a sun gearlubricant stream may be oriented from an oil supply onto the engagementbetween the sun gear and the planet gear.

For an efficient discharge of the oil, the flow guidance device, inparticular the collecting region of the flow guidance device, may befluidically connected to an oil discharge device.

In order to capture as much of the planetary gear lubricant stream aspossible, in one embodiment the flow cross-section of the outlet regionof the collecting region may be between 0.5 mm and 4 mm. In principle,the flow cross-section should be as small as possible. Here, the flowcross-section of the outlet region of the collecting region may bedesigned to be adjustable by an adjustment device, so that the oilsupply can be adapted to different operating states.

For example, the angle of the edge relative to the planet gear, and/orthe distance from the planet gear, may be adjustable by the adjustmentdevice.

Furthermore, in one embodiment, the flow guidance device may have atleast partially a concave region which is open to the collecting regionand in particular is part of the collecting region.

Such a planetary gear box may be used in a gas turbine engine for anaircraft.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine, for example an aircraft engine. Such a gas turbineengine may comprise a core engine comprising a turbine, a combustor, acompressor, and a core shaft connecting the turbine to the compressor.Such a gas turbine engine may comprise a fan (with fan blades) which ispositioned upstream of the core engine.

Arrangements of the present disclosure may be advantageous inparticular, but not exclusively, for geared fans, which are driven via agear box. Accordingly, the gas turbine engine may comprise a gear boxwhich is driven via the core shaft and whose output drives the fan insuch a way that it has a lower rotational speed than the core shaft. Theinput to the gear box may be provided directly from the core shaft, orindirectly via the core shaft, for example via a spur shaft and/or spurgear. The core shaft may be connected rigidly to the turbine and thecompressor, such that the turbine and compressor rotate at the samerotational speed (with the fan rotating at a lower rotational speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The core engine mayfurthermore comprise a second turbine, a second compressor, and a secondcore shaft, which connects the second turbine to the second compressor.The second turbine, the second compressor and the second core shaft maybe arranged so as to rotate at a higher rotational speed than the firstcore shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) a flow from the first compressor.

The gear box may be designed to be driven by the core shaft that isconfigured to rotate (for example during use) at the lowest rotationalspeed (for example the first core shaft in the example above). Forexample, the gear box may be designed to be driven only by the coreshaft that is configured to rotate (for example during use) at thelowest rotational speed (for example only by the first core shaft andnot by the second core shaft, in the example above). Alternatively, thegear box may be designed to be driven by one or more shafts, for examplethe first and/or second shaft in the example above.

In a gas turbine engine as described and/or claimed herein, a combustormay be provided axially downstream of the fan and compressor (orcompressors). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, if a secondcompressor is provided. By way of a further example, the flow at theexit of the compressor may be fed to the inlet of the second turbine,when a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and the secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a series of rotorblades and a series of stator blades, which may be variable statorblades (that is to say the angle of attack may be variable). The seriesof rotor blades and the series of stator blades may be axially offsetfrom one another.

The or each turbine (for example the first turbine and the secondturbine as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator blades. The series of rotor blades and the series ofstator blades may be axially offset from one another.

Each fan blade may have a radial span extending from a root (or a hub)at a radially inner location over which gas flows, or from a spanposition of 0%, to a tip at a span position of 100%. The ratio of theradius of the fan blade at the hub to the radius of the fan blade at thetip may be less than (or of the order of) any of the following: 0.4,0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28,0.27, 0.26 or 0.25. The ratio of the radius of the fan blade at the hubto the radius of the fan blade at the tip may be in a closed intervaldelimited by two values in the previous sentence (that is to say thevalues may form upper or lower limits). These ratios can commonly bereferred to as the hub-to-tip ratio. The radius at the hub and theradius at the tip may both be measured at the leading edge (or theaxially forwardmost edge) of the blade. The hub-to-tip ratio refers, ofcourse, to that portion of the fan blade over which gas flows, i.e. theportion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of the fan blade at its leading edge. The diameter of the fan(which can generally be double the radius of the fan) may be larger than(or of the order of): 250 cm (approximately 100 inches), 260 cm, 270 cm(approximately 105 inches), 280 cm (approximately 110 inches), 290 cm(approximately 115 inches), 300 cm (approximately 120 inches), 310 cm,320 cm (approximately 125 inches), 330 cm (approximately 130 inches),340 cm (approximately 135 inches), 350 cm, 360 cm (approximately 140inches), 370 cm (approximately 145 inches), 380 cm (approximately 150inches), or 390 cm (approximately 155 inches). The fan diameter may bein an inclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in operation. Generally, therotational speed is lower for fans with a larger diameter. Purely as anon-limiting example, the rotational speed of the fan under cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of a further non-limiting example, the rotational speed ofthe fan under cruise conditions for an engine having a fan diameter inthe range of from 250 cm to 300 cm (for example 250 cm to 280 cm) may bein the range of from 1700 rpm to 2500 rpm, for example in the range offrom 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to2100 rpm. Purely by way of a further non-limiting example, therotational speed of the fan under cruise conditions for an engine havinga fan diameter in the range of from 320 cm to 380 cm may be in the rangeof from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpmto 1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

During the use of the gas turbine engine, the fan (with associated fanblades) rotates about an axis of rotation. This rotation results in thetip of the fan blade moving with a speed U_(tip). The work done by thefan blades on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the average 1-D enthalpy rise) across the fan andU_(tip) is the (translational) speed of the fan tip, for example at theleading edge of the tip (which may be defined as fan tip radius at theleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be more than (or of the order of): 0.3, 0.31, 0.32, 0.33,0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (wherein all units in thispassage are Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure can haveany desired bypass ratio, wherein the bypass ratio is defined as theratio of the mass flow rate of the flow through the bypass duct to themass flow rate of the flow through the core at cruise conditions. In thecase of some arrangements, the bypass ratio can be more than (or of theorder of): 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5,16, 16.5, or 17. The bypass ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds). The bypass duct may be substantiallyannular. The bypass duct may be situated radially outside the coreengine. The radially outer surface of the bypass duct may be defined byan engine nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the ram pressure upstreamof the fan to the ram pressure at the exit of the highest pressurecompressor (before entry into the combustor). By way of a non-limitingexample, the overall pressure ratio of a gas turbine engine as describedand/or claimed herein at constant speed can be greater than (or in themagnitude of): 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressureratio may be in an inclusive range bounded by any two of the values inthe previous sentence (i.e. the values may form upper or lower bounds).

The specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. The specificthrust of an engine as described and/or claimed herein at cruiseconditions may be less than (or of the order of magnitude of): 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines can be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of a non-limiting example, a gasturbine as described and/or claimed herein may be capable of generatinga maximum thrust of at least (or of the order of magnitude of): 160 kN,170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN,500 kN or 550 kN. The maximum thrust may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The thrust referred to above maybe the maximum net thrust under standard atmospheric conditions at sealevel plus 15° C. (ambient pressure 101.3 kPa, temperature 30° C.), withthe engine static.

During use, the temperature of the flow at the entry to thehigh-pressure turbine can be particularly high. This temperature, whichmay be referred to as TET, may be measured at the exit to the combustor,for example directly upstream of the first turbine blade, which in turnmay be referred to as a nozzle guide blade. At cruising speed, the TETmay be at least (or of the order of magnitude of): 1400 K, 1450 K, 1500K, 1550 K, 1600 K or 1650 K. The TET at cruise may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The maximum TET in the useof the engine may be at least (or of the order of magnitude of), forexample: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, under a high thrustcondition, for example under a maximum take-off thrust (MTO) condition.

A fan blade and/or airfoil portion of a fan blade described and/orclaimed herein may be produced from any suitable material or combinationof materials. For example at least a part of the fan blade and/orairfoil may be produced at least in part from a composite, for example ametal matrix composite and/or an organic matrix composite, such ascarbon fibre. As a further example, at least a part of the fan bladeand/or aerofoil may be produced at least in part from a metal, such ase.g. a titanium-based metal or an aluminium-based material (such as e.g.an aluminium-lithium alloy) or a steel-based material. The fan blade maycomprise at least two regions produced using different materials. Forexample, the fan blade may have a protective leading edge, which isproduced using a material that is better able to resist impact (forexample from birds, ice or other material) than the rest of the blade.Such a leading edge may, for example, be produced using titanium or atitanium-based alloy. Thus, purely by way of example, the fan blade mayhave a carbon-fibre-based or aluminium-based body (such as analuminium-lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture device whichcan engage with a corresponding slot in the hub (or disk). Purely as anexample, such a fixture may be in the form of a dovetail that may slotinto and/or be brought into engagement with a corresponding slot in thehub/disk in order to fix the fan blade to the hub/disk. By way offurther example, the fan blades may be formed integrally with a centralportion. Such an arrangement may be referred to as a blisk or a bling.Any arbitrary suitable method may be used for production of such a bliskor bling. For example, at least a part of the fan blades may be machinedfrom a block and/or at least part of the fan blades may be attached tothe hub/disk by welding, such as for example linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle can allow the exit cross section of the bypass duct to be variedduring operation. The general principles of the present disclosure canapply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean the cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions can be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or the engine between (in terms of time and/or distance) the top ofclimb and the start of descent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example of the order of Mach0.8, of the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anyarbitrary speed within these ranges can be the constant cruisecondition. In the case of some aircraft, the constant cruise conditionsmay be outside these ranges, for example below Mach 0.7 or above Mach0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example of the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to thefollowing: a forward Mach number of 0.8, a pressure of 23000 Pa and atemperature of −55° C.

As used anywhere herein, “cruising speed” or “cruise conditions” maymean the aerodynamic design point. Such an aerodynamic design point (orADP) may correspond to the conditions (including, for example, the Machnumber, ambient conditions and thrust requirement) for which the fanoperation is designed. This may mean, for example, the conditions underwhich the fan (or gas turbine engine) has the optimum efficiency interms of construction.

During operation, a gas turbine engine described and/or claimed hereinmay be operated under the cruise conditions defined elsewhere herein.Such cruise conditions may be determined by the cruise conditions (forexample the conditions during the middle part of the flight) of anaircraft to which at least one (for example two or four) gas turbineengine(s) can be fastened in order to provide propulsive thrust.

It is self-evident to a person skilled in the art that a feature orparameter described in relation to one of the above aspects may beapplied to any other aspect, unless these are mutually exclusive.Furthermore, any feature or any parameter described here may be appliedto any aspect and/or combined with any other feature or parameterdescribed here, unless these are mutually exclusive.

Embodiments will now be described by way of example with reference tothe figures, in which:

FIG. 1 a sectional side view of a gas turbine engine;

FIG. 2 shows a close-up sectional side view of an upstream portion of agas turbine engine;

FIG. 3 shows a partially cut-away view of a gear mechanism for a gasturbine engine;

FIG. 4 shows a sectional view through a part of an embodiment of aplanetary gear box with a flow guidance means for a planet gearlubricant oil stream;

FIG. 5A shows a schematic detail view of a first embodiment with an edgefor controlling the planet gear lubricant stream;

FIG. 5B shows a schematic detail view of a second embodiment with anedge for controlling the planet gear lubricant stream;

FIG. 6 shows a schematic illustration of an embodiment with anadjustable flow guidance device;

FIG. 7 shows a schematic illustration of an embodiment with a flowguidance device with portions opened to the collecting region.

FIG. 1 illustrates a gas turbine engine 10 having a main axis ofrotation 9. The engine 10 comprises an air intake 12 and a fan 23 thatgenerates two air flows: a core air flow A and a bypass air flow B. Thegas turbine engine 10 comprises a core 11 that receives the core airflow A. When viewed in the order corresponding to the axial direction offlow, the core engine 11 comprises a low-pressure compressor 14, ahigh-pressure compressor 15, a combustion device 16, a high-pressureturbine 17, a low-pressure turbine 19, and a core thrust nozzle 20. Anengine nacelle 21 surrounds the gas turbine engine 10 and defines abypass duct 22 and a bypass thrust nozzle 18. The bypass air flow Bflows through the bypass duct 22. The fan 23 is attached to and drivenby the low-pressure turbine 19 via a shaft 26 and an epicyclic planetarygear box 30.

During operation, the core air flow A is accelerated and compressed bythe low-pressure compressor 14 and directed into the high-pressurecompressor 15, where further compression takes place. The compressed airexpelled from the high-pressure compressor 15 is directed into thecombustion device 16, where it is mixed with fuel and the mixture iscombusted. The resulting hot combustion products then propagate throughthe high-pressure and the low-pressure turbines 17, 19 and thereby drivesaid turbines, before being expelled through the nozzle 20 to provide acertain propulsive thrust. The high-pressure turbine 17 drives thehigh-pressure compressor 15 by way of a suitable connecting shaft 27.The fan 23 generally provides the major part of the propulsive thrust.The epicyclic planetary gear box 30 is a reduction gear box.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low-pressure turbine 19 (see FIG. 1 ) drives the shaft26, which is coupled to a sun wheel 28 of the epicyclic planetary gearmechanism 30. Multiple planet gears 32, which are coupled to one anotherby a planet carrier 34, are situated radially to the outside of the sungear 28 and mesh therewith. The planet carrier 34 guides the planetgears 32 in such a way that they circulate synchronously around the sungear 28, whilst enabling each planet gear 32 to rotate about its ownaxis. The planet carrier 34 is coupled via linkages 36 to the fan 23 inorder to drive its rotation about the engine axis 9. An external gear orring gear 38 that is coupled via linkages 40 to a stationary supportstructure 24 is situated radially to the outside of the planet gears 32and meshes therewith.

It should be noted that the expressions “low-pressure turbine” and“low-pressure compressor”, as used herein, can be taken to mean thelowest-pressure turbine stage and lowest-pressure compressor stage (i.e.not including the fan 23), respectively, and/or the turbine andcompressor stages that are connected together by the connecting shaft 26with the lowest rotational speed in the engine (i.e. not including thegear box output shaft that drives the fan 23). In some documents, the“low-pressure turbine” and the “low-pressure compressor” referred toherein may alternatively be known as the “intermediate-pressure turbine”and “intermediate-pressure compressor”. Where such alternativenomenclature is used, the fan 23 may be referred to as a first, orlowest-pressure, compression stage.

The epicyclic planetary gear box 30 is shown in greater detail by way ofexample in FIG. 3 . The sun gear 28, planet gears 32 and ring gear 38 ineach case comprise teeth on their periphery to allow meshing with theother toothed gears. However, for clarity, only exemplary portions ofthe teeth are illustrated in FIG. 3 . Although four planet gears 32 areillustrated, it will be apparent to a person skilled in the art thatmore or fewer planet gears 32 may be provided within the scope ofprotection of the claimed invention. Practical applications of anepicyclic planetary gear box 30 generally comprise at least three planetgears 32.

The epicyclic planetary gear box 30 illustrated by way of example inFIGS. 2 and 3 is a planetary gear box in which the planet carrier 34 iscoupled to an output shaft via linkages 36, with the ring gear 38 beingfixed. However, any other suitable type of planetary gear box 30 may beused. As a further example, the planetary gear box 30 may be a stararrangement, in which the planet carrier 34 is held fixed, with the ringgear (or external gear) 38 being allowed to rotate. In such anarrangement, the fan 23 is driven by the ring gear 38. As a furtheralternative example, the gear box 30 may be a differential gear box inwhich both the ring gear 38 and the planet carrier 34 are allowed torotate.

It is self-evident that the arrangement shown in FIGS. 2 and 3 is merelyan example, and various alternatives fall within the scope of protectionof the present disclosure. Purely by way of example, any suitablearrangement can be used for positioning the gear box 30 in the engine 10and/or for connecting the gear box 30 to the engine 10. By way of afurther example, the connections (such as the linkages 36, 40 in theexample of FIG. 2 ) between the gear box 30 and other parts of theengine 10 (such as the input shaft 26, the output shaft and the fixedstructure 24) may have a certain degree of stiffness or flexibility. Asa further example, any suitable arrangement of the bearings betweenrotating and stationary parts of the engine 10 (for example between theinput and output shafts of the gear box and the fixed structures, suchas the gear casing) may be used, and the disclosure is not limited tothe exemplary arrangement of FIG. 2 . For example, where the gear box 30has a star arrangement (described above), the person skilled in the artwould readily understand that the arrangement of output and supportlinkages and bearing positions would typically be different to thatshown by way of example in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gear box types (for example star-shaped orepicyclic-planetary), support structures, input and output shaftarrangement, and bearing positions.

Optionally, the gear box may drive additional and/or alternativecomponents (for example the intermediate-pressure compressor and/or abooster compressor).

Other gas turbine engines in which the present disclosure can be usedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of connecting shafts. By way of a further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22,meaning that the flow through the bypass duct 22 has its own nozzle thatis separate from and radially outside the core engine nozzle 20.However, this is not restrictive, and any aspect of the presentdisclosure can also apply to engines in which the flow through thebypass duct 22 and the flow through the core 11 are mixed or combinedbefore (or upstream of) a single nozzle, which may be referred to as amixed flow nozzle. One or both nozzles (whether mixed or split flow) canhave a fixed or variable region. Although the example described relatesto a turbofan engine, the disclosure may be applied for example to anytype of gas turbine engine, for example an open-rotor engine (in whichthe fan stage is not surrounded by an engine nacelle) or a turbopropengine.

The geometry of the gas turbine engine 10, and components thereof,is/are defined by a conventional axis system, which comprises an axialdirection (which is aligned with the axis of rotation 9), a radialdirection (in the direction from bottom to top in FIG. 1 ), and acircumferential direction (perpendicular to the view in FIG. 1 ). Theaxial, radial and circumferential directions are mutually perpendicular.

The planetary gear box 30 of the gas turbine engine 10 must, inparticular in turbofan engines, function for a very long time withoutmaintenance. Therefore an efficient and robust lubrication, here withoil, is of particular importance. The flow of lubricating oil in theinterior of the planetary gear box 30 must take into account the ageingof the oil. The examples illustrated here concern planetary gear boxes30 in which planet gears 34 are driven by a sun gear 28. The outputtakes place via the planet carrier 34. The ring gear 38 is here static.In principle however, other planet gear configurations may also be used.

The description below illustrates embodiments in which a flow guidancemeans 50 serves to guide a planet gear lubricant stream S (i.e. here anoil stream). For visualisation, oil droplets are depicted as small blackdots.

For lubrication, via an oil supply 55, a sun gear lubricant stream S1 isdirected in targeted fashion at the engagement of the sun gear 28 andplanet gears 32. The sun gear 28 here rotates clockwise, and the planetgears 32 accordingly rotate counter-clockwise.

FIG. 4 shows the main flow direction of the sun gear lubricant stream S1as an arrow which is directed at the engagement of the planet gear 32and sun gear 28.

Furthermore, the embodiment of the planetary gear box 30 has an oilsupply 56 for a planet gear lubricant stream S, the main flow directionof which is oriented at least partially along the tooth surfaces and/oralong the periphery of the planet gear 32 after the engagement with thesun gear 28. The oil, drawn in by the engagement between the planetgears 32 and sun gear 28, is then situated between the tooth flanks ofthe planet gears 32 and sun gear 28, and is here subjected to acentrifugal force and transported radially outward.

Because of a blocking of the oil supply by the planet gear lubricantstream S (cooling oil spray 56) and the centrifugal force from therotation of the planet carrier 34, and because the planet gear 32 trailsthe outflung oil in the same direction (rotation of the planet carrier32), an outwardly directed oil stream S is created which follows thecontour of the planet gear 32 (centrifugal force from the rotatingplanet carrier 34). The trajectory of the planet gear lubricant stream Sis formed in particular by the blockade of the lubricant jet at the exitfrom the engagement between the planet gear 32 and sun gear 28.

As shown below, an additional advanced edge 53 of a flow guidance device50 can guide the planet gear lubricant stream S more effectively into acollecting region 51 of the oil guidance component 50.

The main flow direction of this planet gear lubricant stream S isinitially oriented substantially tangentially to the tooth face of theplanet gear 32. The rotation and flow inside the gear casing expand theplanet gear lubricant stream S (illustrated as dots). This can be seenfor example from the formation of a cloud of dots.

The flow guidance device 50, the geometric formation of which isdescribed below, serves to form the planet gear lubricant stream S andconduct this partially out of the gear region in order to avoidsecondary losses. These would occur if the oil of the planet gearlubricant stream S were conveyed further by the planet gear 32 in thedirection of another planet gear 32 or in the direction of the ring gear38. Because of the rotation of the planet carrier 34, a radially outwardflow would then occur.

In order to prevent or at least reduce this, the flow guidance device 50has a collecting region 51 in which the planet gear lubricant stream Sbetween a planet gear 32 and the flow guidance device 50 can becaptured. The collecting region 51 thus prevents an uncontrolledpropagation of the planet gear lubricant stream S into the gear casing.Furthermore, the collecting region 51 has an outlet region 52 which isarranged between the flow guidance device 50 and the tooth surface ofthe planet gear 32. At least a part of the planet gear lubricant streamS can flow out here between the flow guidance device 50 and along thetooth surfaces and/or the periphery of the planet gear 32.

In particular to form this outflow, an edge 53 is arranged in the regionof the outlet region 52 for defined constriction of the passage for theplanet gear lubricant stream S. The edge 53 is made sharp, so it is notdesigned as a type of nozzle which conducts the planet gear lubricantstream S.

The edge 53 shears off a part of the planet gear lubricant stream S andretains it in the collecting region 51. From the collecting region 51,the lubricating oil can be discharged from the immediate area of thetooth engagements via an oil discharge device 54, so that secondarylosses at the planet gears and ring gear, and oil ageing, are reduced.

The edge 54 should be made sharp relative to the planet gear 32, i.e. itis formed in particular at least at a right angle W to the planet gear32 (see e.g. FIGS. 5A, 5B) when the edge is projected onto a tangent atthe planet gear 32 (see FIG. 5A). The edge 54 is shown as a dot in FIG.4 .

The angle W (measured in the rotational direction of the schematicallyillustrated planet gear 32 and again projected onto a tangent to theplanet gear 32) may also be greater than 90°, in particular however lessthe 180° (see FIG. 5B). This ensures that the planet gear lubricantstream S meets a sharp edge when flowing out of the collecting region51.

The flow cross-section of the outlet region 52 of the collecting region51, i.e. the gap width which is delimited on one side by the edge 53,may in typical applications lie between 0.5 and 4 mm. In principle, thisgap should be as small as possible.

For an efficient flow guidance, the main flow of the planet gearlubricant stream S is oriented substantially (e.g. +/−10°) onto the edge54, the outlet region 52 or the collecting region 51, i.e. the planetgear lubricant stream S leaves the oil supply 56 in this direction.

As a whole, the introduction of the edge 53 brings a significantincrease in the captured proportion of the planet gear lubricant streamS. This reduces the ageing of the oil, the temperature of the oil andlosses at the gear box. The reduction in contacts means that less oil isrequired, so that weight can be saved. A fuel saving may lie in theregion of 0.05 to 0.1%.

FIG. 6 shows an embodiment in which an adjustment device 57 may adjust adistance D between the edge 53 of the flow guidance device 50 and theplanet gear 32 (here illustrated schematically), in order to control theplanet gear lubricant stream S. In addition or alternatively, the angleW may also be adjusted by the adjustment device 57.

FIG. 7 shows an embodiment of a flow guidance device 50 in a 3Dillustration. This view corresponds approximately to the side view ofFIG. 4 in the axial direction. However, in FIG. 7 , a part of an oillevel O is marked which lies inside the gear casing (illustrated here).

In the illustration of FIG. 7 , the planet gear lubricant stream Senters from below and is delimited by the sharp edge 53 as describedabove, wherein the planet gear lubricant stream S then flows along theplanet gear (not shown here) in a similar fashion to the embodiment inFIG. 4 .

The interior of the flow guidance device 50 (here shown opened) is hereformed partially as a blade or channel so that the oil level O in theinterior is visible.

A section through the flow guidance device 50 approximately in theradial direction would show a U-shaped profile. The concave side of theblade-like or channel-like part of the flow guidance device 50 is opento the collecting region 51 (here from above and to the left) and isthus a part of the collecting region 51. Here, the channel-likeformation is arranged on the upper part and on the vertical part, sothat oil can sit there.

Some of the oil is then discharged via the oil discharge device 54.

It goes without saying that the invention is not limited to theembodiments described above, and various modifications and improvementscan be made without departing from the concepts described herein. Any ofthe features may be used separately or in combination with any otherfeatures, unless they are mutually exclusive, and the disclosure extendsto and includes all combinations and subcombinations of one or morefeatures that are described herein.

LIST OF REFERENCE SIGNS

-   -   9 Main rotation axis    -   10 Gas turbine engine    -   11 Core engine    -   12 Air inlet    -   14 Low-pressure compressor    -   15 High-pressure compressor    -   16 Combustion device    -   17 High-pressure turbine    -   18 Bypass thrust nozzle    -   19 Low-pressure turbine    -   20 Core thrust nozzle    -   21 Engine nacelle    -   22 Bypass duct    -   23 Fan    -   24 Stationary support structure    -   26 Shaft    -   27 Connecting shaft    -   28 Sun gear    -   30 Gear box, planetary gear box    -   32 Planet gears    -   34 Planet carrier    -   36 Linkage    -   38 Ring gear    -   40 Linkage    -   50 Flow guide device    -   51 Collecting region    -   52 Outlet region of collecting region    -   53 Edge    -   54 Oil discharge device    -   55 Oil supply for sun gear lubricant stream S1    -   56 Oil supply for planet gear lubricant stream S    -   A Core air flow    -   B Bypass air flow    -   O Oil level    -   S Planet gear lubricant stream    -   S1 Sun gear lubricant stream    -   W Angle

The invention claimed is:
 1. A planetary gear box, comprising: a flowguidance device for a planet gear lubricant stream in the planetary gearbox, which is directed at least partially along a planet gear, after anengagement with a sun gear, a collecting region of the flow guidancedevice in which the planet gear lubricant stream between the planet gearand the flow guidance device is at least partially captured, wherein thecollecting region has an outlet region between the flow guidance deviceand a tooth face of the planet gear for at least a part of the planetgear lubricant stream, and an edge arranged at the outlet region fordefined constriction of a passage for the planet gear lubricant stream,wherein the edge is a sharp edge having an included angle between facesof less than 90° and is angled backward toward a rotational direction ofthe tooth face of the planet gear.
 2. The planetary gear box accordingto claim 1, wherein an angle, measured in a rotational direction of theplanet gear between the edge and a tangent to the planet gear, isgreater than 90°, and less than 180°.
 3. The planetary gear boxaccording to claim 1, wherein a main flow direction of the planet gearlubricant stream is oriented from an oil supply onto the collectingregion of the flow guidance device, the edge or the outlet region of theflow guidance device.
 4. The planetary gear box according to claim 3,wherein an orientation of the main flow direction of the planet gearlubricant stream deviates by +/−20° from an exact linear geometricorientation.
 5. The planetary gear box according to claim 1, wherein asun gear lubricant stream is oriented from an oil supply onto anengagement between the sun gear and the planet gear.
 6. The planetarygear box according to claim 1, wherein the collecting region of the flowguidance device is fluidically connected to an oil discharge device. 7.The planetary gear box according to claim 1, wherein a flowcross-section of the outlet region of the collecting region is between0.5 mm and 4 mm.
 8. The planetary gear box according to claim 1, andfurther comprising an adjustment device, wherein a flow cross-section ofthe outlet region of the collecting region is adjustable by theadjustment device.
 9. The planetary gear box according to claim 8,wherein the adjustment device is configured to adjust at least onechosen from an angle of the edge relative to the planet gear and adistance between the flow guidance device and the planet gear.
 10. Theplanetary gear box according to claim 1, wherein the flow guidancedevice has at least partially a concave region which is open to thecollecting region and is part of the collecting region.
 11. A gasturbine engine for an aircraft, comprising: a core engine comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan, which is positioned upstream of the core engine,wherein the fan comprises a plurality of fan blades; and a gear box,which can be driven by the core shaft, wherein the fan can be driven bymeans of the gear box at a lower speed than the core shaft, wherein thegear box is the planetary gear box according to claim
 1. 12. Theplanetary gear box according to claim 1, wherein an angle, measured in arotational direction of the planet gear between the edge and a tangentto the planet gear, is greater than 90° and less than 120°.
 13. Theplanetary gear box according to claim 3, wherein an orientation of themain flow direction of the planet gear lubricant stream deviates by+/−10° from an exact linear geometric orientation.